1. Field of the Invention
The present invention relates generally to improving the metallurgical fatigue life of certain general aviation aircraft airframe structures, and more particularly, but not by way of limitation, to enhancement and extension of the structural integrity thereof.
2. Discussion
It is known that aluminum aircraft structures accumulate metal fatigue with repetitive cyclic loading imposed on the structures as a result of flight loads and loading imposed from takeoffs and landings.
For various historic, economic and market reasons, the average age in flight hours of the general aviation fleet is increasing each year. Because the accumulation of fatigue in the aircraft metal structure tends to follow the flight hours, there is an ever increasing risk of fatigue accumulating to the point that many general aviation aircraft will have to be grounded because of unacceptable risks of fatigue cracks in primary airframe structures.
When sufficient fatigue damage has accumulated in the airframe, the accumulated fatigue will cause small cracks to initiate in localized portions of the airframe structure. As fatigue continues to accumulate, the cracks will inevitably grow, and if undetected and not repaired, there will eventually be failure of the airframe and probable loss of life if in flight.
It is known that fatigue damage is dramatically accelerated with increasing-load levels during repetitive cyclic loading of the airframes. This cumulative damage is highly nonlinear. Even small increases in loading levels result in disproportionately shortened fatigue life for airframe structures. Conversely, if critical portions of the structure can be identified and the loading on the critical structures reduced even by modest proportions, the useful fatigue life of the airframe can be greatly increased.
The fleet of Beechcraft Bonanza, Baron and T-34 airplanes manufactured by Beech Aircraft Corporation, now Raytheon Aircraft Corporation, all share a common design arrangement in the main wing structure. This common design arrangement consists of, among many other features, wing members that can be independently removed from the fuselage. The wings are attached to the fuselage structure with four bolts on each side, of which two are part of the front spar and two are part of the rear wing spar attachment structure on each side of the fuselage structure.
Within the fuselage is located a crucial portion of the total main wing spar structure, which is commonly referred to as the center spar carry through structure, or simply, for the purpose of this disclosure, the center spar. This structure is designed to tie the two wing spars together as an integral unit and to transfer the load from the two wings through the fuselage.
The center spar assembly consists of upper and lower principal structure elements. The upper portion of the center spar structure is normally in compressive loading, while the lower portion of the center spar structure is normally in tensile loading, both of which will vary with the flight and landing-loading.
Among features that commonly appear in these types of structures are the presences of highly loaded tension bolts that present potential failure points in the most critical portions of the structure. These failure modes result because there is no redundant load path to carry loads in the event one of the highly loaded tension bolts fails. Those skilled in the art recognize that avoiding critical structural features characterized by single point failure modes is desirable in the design of structures upon which human safety depends.
The normal loading for the aircraft operating in smooth air imposes tensile loading on the lower spar section proportional to the acceleration that the aircraft is experiencing with respect to the vertical axis through the aircraft structure. This acceleration is commonly referred to as the maneuver loading or ‘g-load’ for the aircraft. During turbulence, and during accidental or deliberate maneuvers, the g-load increases substantially, and depending on the particular type of aircraft, it can reach six times its normal loading for aerobatic maneuvers, and more during unintentional overload events.
In recent years, observations and investigations of accidents and incidents that involved similar aircraft and structures that had experienced long term repetitive high loading conditions revealed the presence of fatigue cracks that resulted in actual or pending structural failure. Other investigations revealed that the original detail design of the center carry through structure resulted in the creation of critical structural stress risers at specific locations that, if not mitigated, result in premature or early retirement of the airframe due to fatigue accumulation. The fatigue damage observed is almost entirely the result of tensile loading in such areas, as opposed to compressive loading.
Mitigation of the stress concentration features in the critical center fuselage structure is exceptionally difficult because access to the structure is restricted by its inherent original design, and disassembly is labor intensive and essentially cost prohibitive.
There have been attempts to mitigate fatigue accumulation of these types of structures by various means, including the installation of a ‘spar strap.’ This is a device in which a steel strap is arranged across the exterior of the belly of the fuselage of the aircraft and attached to the outboard area of each wing. This attempt at mitigation has several undesirable effects. Among these is that the steel and aluminum over large unprotected areas can cause dissimilar metal corrosion. Another is that the modification is labor intensive and expensive to install, and the installation restricts access to the areas of the wings that normally require repetitive maintenance inspections. Yet another large disadvantage of external spar strap arrangements is the additional resultant airframe drag, and the consequent reduction in speed, range and utility of the aircraft.
The present invention provides an effective and cost efficient solution to the problem discussed above while avoiding the disadvantages of the prior art.